The present invention relates to an infrared suppression system (IRSS), and more particularly to an infrared suppression system which uses a film-cooled spiral septum within an exhaust duct to mask the engine exhaust and reduce the overall infrared signature of an aircraft.
The exhaust plume and plume-heated surfaces from a gas turbine engine may be a source of infrared (IR) energy which may be detected by heat seeking missiles and/or various forms of infrared imaging systems for targeting/tracking purposes. With respect to the former, a heat-seeking missile may obtain directional cues from infrared energy generated by the engine exhaust such that the amount of infrared energy is one of the primary determining factors of a missile's accuracy, and consequently, lethality. Regarding the latter, infrared imaging systems detect and amplify infrared energy for detection and/or targeting.
IR suppression systems are utilized on many military aircraft including most rotary wing aircraft to provide IR signature reduction. Future IR threats, however, will require even greater levels of IR signature reduction.
Generally, IR suppression systems are primarily designed to: (a) reduce the infrared energy below a threshold level of a perceived threat; (b) maintain engine performance; and (c) minimize weight and packaging associated therewith. Secondary consequences may include: (i) minimizing system or configuration complexity to reduce fabrication and maintainability costs; and (ii) minimizing the external aerodynamic drag produced by such IR suppressor systems.
Current IR suppression systems for rotary wing aircraft are primarily designed to mix the high temperature exhaust flow with a cooling airflow supplied by a mixing duct which communicates with an engine exhaust duct. The mixing of large amounts of ambient air with the engine exhaust reduces the overall gas temperature prior to “dumping” the engine exhaust overboard thus lowering the aircraft IR signature. To achieve significant reductions in temperature, however, a relatively significant volume of ambient air must mixed into the high temperature exhaust flow. This requires relatively large intakes and a final exhaust stage which provides a flow area capacity for the combined engine exhaust flow volume and the additional mixed in ambient airflow volume.
Such conventional IR suppressor systems are limited by packaging space restrictions thereof in which a relatively significant area is required to provide ample mixing and flow area. Adaptation to relatively small rotary wing aircraft or retrofitting to current aircraft packaging constraints may be limited and appropriate for designs with less stringent packaging space restrictions.
A high IR signature source on an unsuppressed aircraft is heated exhaust components (e.g. power turbine, deswirl vanes, duct walls). As such, one effective method of signature reduction is shadowing, masking, or cooling of these surfaces to minimize infrared emissions over particular viewing angles.
It is also desirable to minimize impingement of hot engine exhaust on adjacent aircraft structure so as to avoid, the generation of a “hot spot” separate from the primary source associated with the nozzle/exhaust plume. Disadvantageously, the mixing operation reduces the velocity of the exhaust flow such that the exhaust velocity may be too low to expel the exhaust far enough from the fuselage to avoid such a “hot spot.” A farther disadvantage is that if the exhaust may not have enough velocity to escape rotor downwash and the exhaust gas may be re-ingested into the engines which reduces engine efficiency.
Accordingly, it is desirable to provide an Infrared Suppression System (IRSS) which reduces the overall IR signature of the aircraft, is compact in design, masks the IR energy emitted/radiated from a gas turbine engine for a given viewing/azimuth angle, and minimizes impingement of engine exhaust onto adjacent aircraft structure while maintaining engine performance and residual thrust.